1. Field of the Invention
The present invention relates generally to fluid reaction surfaces, and more specifically to turbine airfoils with a cooling circuit.
2. Description of the Related Art Including Information Disclosed under 37 CFR 1.97 and 1.98
Turbine airfoils, such as rotor blades and stator vanes, pass cooling air through complex cooling circuits within the airfoil to provide cooling from the extreme heat loads on the airfoil. A gas turbine engine passes a high temperature gas flow through the turbine to produce power. The engine efficiency can be increased by increasing the temperature of the gas flow entering the turbine. Therefore, an increase in the airfoil cooling can result in an increase in engine efficiency.
Prior art airfoil cooling of blades makes use of a single five-pass aft flowing serpentine cooling circuit 11-15 comprised of a forward section leading edge impingement cavity 17 and an aft flowing serpentine flow channels with airfoil trailing edge discharge cooling holes 20 as seen in FIG. 1. In the forward section of the blade leading edge impingement cooling, it is normally designed in conjunction with leading edge backside impingement plus showerhead and pressure side and suction side film discharge cooling holes. Cooling air is supplied from the first up-pass of the 5-pass serpentine flow circuit. The impingement cooling air is normally fed through a row of metering holes 16, impinged onto the backside of the airfoil leading edge surface to provide backside impingement cooking prior to discharging through the three showerhead holes 18 and pressure side and suction side gill holes 19.
In the prior art 5-pass aft flowing serpentine cooling circuit of FIG. 1, the internal cavities are constructed with internal ribs connecting the airfoil pressure and suction walls. In most of the cases, the internal cooling cavities are at low aspect ratios which is subject to high rotational affect on the cooling side heat transfer coefficient. In addition, the low aspect ratio cavity yields a very low internal cooling side convective area ratio to the airfoil hot gas external surface.
The object of the present invention is to provide for a blade with a cooling circuit that provides for a near wall spiral flow cooling arrangement which optimizes the airfoil mass average sectional metal temperature to improve airfoil creep capability for a blade cooling design.
Another object of the present invention is to maximize the airfoil cooling performance for a given amount of cooling air and minimize the Coriolis effects due to rotation on the airfoil internal cavities' heat transfer performance.